For the remainder of the present disclosure, a longitudinal axis corresponds to the axis which extends from the front tip to the rear tip of an aircraft. A longitudinal direction is a direction parallel to the longitudinal axis. A longitudinal plane corresponds to a plan containing the longitudinal axis. A transverse plane corresponds to a plane perpendicular to the longitudinal axis. A radial direction is a direction perpendicular to the longitudinal axis. An internal face or surface corresponds to a face or surface oriented towards the interior of the fuselage and an external face or surface corresponds to a face or surface oriented towards the exterior of the fuselage.
As is known, the fuselage of an aircraft comprises a structure which is principally responsible for taking up the forces and to which is attached an envelope, also called the skin, which gives the aircraft its aerodynamic properties.
According to an embodiment shown in FIGS. 1A and 1B, the skin comprises multiple juxtaposed panels 10, reinforced by stringers 12 attached to the internal face of the panels 10. The stringers 12 are arranged in longitudinal planes.
The structure comprises frames 14 which are arranged in transverse planes and are spaced apart in the longitudinal direction. In longitudinal planes, the frames 14 have a Z-shaped cross section with a web 16 arranged in a transverse plane and two flanges 18, 18′ connected to the web 16 and perpendicular to the web 16.
In order to form the fuselage, the reinforced panels 10 and the frames 14 are connected using clips 20.
Each clip 20 comprises at least two flanges 22.1 and 22.2 connected to one another so as to form an angle bar, a first flange 22.1 clamped against the internal face of a panel 10 and a second flange 22.2 clamped against the web 16 of a frame 14.
As shown in FIG. 1A, a clip 20′ may have a more complex shape and may comprise a bracket 24 connecting two flanges.
The first flange 22.1 of each clip 20 is secured to the panel 10 by attachment elements 26 such as bolts or rivets. These attachment elements 26 comprise milled heads which are housed in milled portions provided in the external face of the panel 10 so as to not impair the aerodynamic properties of the fuselage.
The second flange 22.2 of each clip 20 is secured to the web 16 using attachment elements 28 such as bolts or rivets.
According to a first operating method, each panel 10 is produced and reinforced with the stringers 12 while flat. This solution makes it possible to simplify the method for manufacturing the reinforced panels 10. Thus, in the case of panels 10 and stringers 12 made of composite material, it is possible to automate the method for manufacturing the reinforced panels, in particular by using automatic lay-up machines.
Then, the planar panel 10 reinforced with the stringers 12 is shaped on a first tool so as to give it its curved shape. This curved shape varies depending on the position of the panel 10 in the fuselage. After this shaping step, the external face of the panel 10 extends along a reference surface.
Depending on the position of the panel in the fuselage, this reference surface is developable and approximately semi-cylindrical or it is not developable for the panels close to the front tip or to the rear tip of the aeroplane.
This first tool comprises a scaffold which supports at least one support surface against which is clamped the internal face of the panel 10. The scaffold and the support surfaces are designed to be rigid and to retain their shape when the panel deforms.
The frames 14 are arranged one by one on a second tool. This second tool makes it possible to position the frames 14 according to their reference positions in the fuselage.
This second tool comprises a scaffold which supports ring portions called ribs, one for each frame. Each rib comprises a support surface against which is clamped the web of the frame in order to position it in a transverse plane and three locating pins for positioning and immobilizing the frame in the transverse plane.
Once the frames are positioned, the panel 10, reinforced and shaped along the reference surface, is positioned relative to the frames 14.
After positioning the panel 10 relative to the frames 14, the clips 20, the frames and the panel are assembled. For each clip 20, a first flange 22.1 of the clip comes to bear against the internal face and a second flange 22.2 of the clip comes to bear against the web 16 of the frame. The clips are positioned so as to be distributed all along each frame. In order to facilitate this placement, the second tool comprises, for each clip, a stop by which it can be positioned along the frame. Before the attachment elements 26, 28 are installed, the clips are held in position using pins.
In order to control the contact force between the assembled parts, a clamping force is applied to the parts to be assembled at every tenth attachment point. The clearance between the parts to be assembled must be of the order of 0.3 mm for a force of 20 daN. If this clearance is greater than 0.3 mm, shims are arranged between the parts. This method consisting in measuring, under load, the clearance between the parts makes it possible to control the internal stresses induced during assembly.
Finally, the attachment elements 26, 28 are installed so as to connect the clip 20 to the panel 10 on one hand, and the clip 20 to the frame 14 on the other hand.
This first operating method is difficult to automate. A first constraint is that it is necessary to provide a first automatic installation means arranged inside the fuselage in order to install the attachment elements 28 connecting the clips to the frames, and a second automatic installation means outside the fuselage in order to install the attachment elements 26 between the panel and the clips which must be installed from outside the fuselage because of the presence of the milled portions. Another constraint is that the second tool must make it possible to hold the elements to be assembled, that is to say the panel, the frames and the clips according to their reference positions. Given these constraints, this assembly is carried out by operators equipped with drilling units.
According to a second operating method described in document EP-2.404.824, the frames are arranged one by one on a tool according to their reference positions in the fuselage.
To that end, the tool comprises a scaffold which supports ring portions called ribs, one for each frame. Each rib comprises a support surface against which is clamped the web of the frame in order to position it in a transverse plane and three locating pins in order to position the frame in the transverse plane.
Once the frames are positioned, the flat panel 10, reinforced by the stringers, is deformed on the frames which replace the first tool of the first operating method. Then, the panel and the frames are assembled using connection elements. According to this document, the frames have a particular hollow cross section, and each comprise a face clamped against the panel such that it is possible to dispense with the clips.
Even though, in theory, this second operating method can make it possible to automate the assembly using an automatic means for installing attachment elements from outside the fuselage, it is not satisfactory for the following reasons.
A first drawback is that this operating method does not make it possible to control the internal stresses induced during assembly of the panel and of the frames, in particular if the panel has a reference surface which is not developable.
At the contact points between the frame and the panel, it is not possible to determine the contact force between the frame and the panel and thus control the stresses induced during assembly.
This is also the case for those regions in which the clearance is less than 0.3 mm when under load.
If the clearance between the frame and the panel when under load is greater than 0.3 mm, it is necessary to put in place shims. It is then necessary to remove the panel to put in place the shims then reposition it once the shims have been put in place. However, in the case of a panel with a reference surface which is not developable, it is very difficult to reposition it in the same manner.
Another drawback is that, when the contact surfaces of the parts to be assembled are coated with a sealing mastic, this mastic prevents the parts from sliding relative to one another in order to adjust their relative position.
Finally, another drawback is that the frames must be rigid in order not to deform when the panel is shaped. Consequently, frames having a Z-shaped cross section cannot be used as this cross section does not provide sufficient torsional and flexural inertia. In order to stiffen them, it would be possible to increase the number of locating pins used to position each frame on its rib. However, in this case, as for the hollow cross section frames, it is not possible to control the internal stresses induced during assembly.